1. Field of the Invention
The present invention relates generally to a gas turbine engine, and more specifically to an industrial turbine blade with platform cooling.
2. Description of the Related Art Including Information Disclosed Under 37 CFR 1.97 and 1.98
In a gas turbine engine, such as a large frame heavy-duty industrial gas turbine (IGT) engine, a hot gas stream generated in a combustor is passed through a turbine to produce mechanical work. The turbine includes one or more rows or stages of stator vanes and rotor blades that react with the hot gas stream in a progressively decreasing temperature. The efficiency of the turbine—and therefore the engine—can be increased by passing a higher temperature gas stream into the turbine. However, the turbine inlet temperature is limited to the material properties of the turbine, especially the first stage vanes and blades, and an amount of cooling capability for these first stage airfoils.
The first stage rotor blade and stator vanes are exposed to the highest gas stream temperatures, with the temperature gradually decreasing as the gas stream passes through the turbine stages. The first and second stage airfoils (blades and vanes) must be cooled by passing cooling air through internal cooling passages and discharging the cooling air through film cooling holes to provide a blanket layer of cooling air to protect the hot metal surface from the hot gas stream.
FIG. 1 shows a prior art first stage turbine rotor blade used in a large frame heavy duty industrial gas turbine engine. Cooling of the blade platform 12 is produced by passing cooling air through straight cooling channels that have a long length-to-diameter ratio. The pressure side of the platform 12 is cooled with three straight channels 13 each supplied with cooling air through inlet holes 16 that open on the bottom surface of the platform 12 with cooling air from the dead rim cavity located below the platform 12. The suction side of the platform 12 is cooled with three straight channels 15 that are all connected to a larger diameter and longer channel 14 located along the side edge of the platform 12. An inlet hole 16 also supplies the suction side channels 15 with cooling air from the rim cavity. An airfoil 11 extends from the platform 12.
The platform cooling circuit of the FIG. 1 blade suffers from several design problems. Using a film cooling method for the entire blade platform requires a cooling air supply pressure at the dead rim cavity to be at a higher pressure than the peak blade platform external gas side pressure. This platform cooling design induces a high leakage flow around the blade attachment section and therefore causes a performance penalty.
Also, uses long length-to-diameter ration cooling channels that are drilled from the platform edge to the airfoil cooling core in the blade platform wall will produce very high stress levels at the airfoil cooling core and platform cooling channel interface locations that will cause a low blade life. This affect is mainly due to the large mass at the front and back ends of the blade attachment which will constrain any blade platform expansion. Also, with the cooling channels oriented transverse to the primary direction of the stress field, high stress concentrations will occur at the cooling channel inlet holes.
An analysis of the FIG. 1 prior art turbine blade indicates that an over-temperature occurs at the platform pressure side location and at the aft portion of the suction side platform edge and the aft section of the suction side to platform junction.